Fan-turbine rotor assembly with integral inducer section for a tip turbine engine

ABSTRACT

A fan-turbine rotor assembly for a tip turbine engine includes a fan hub with an outer periphery scalloped by a multitude of elongated openings which extend into a fan hub web. Each elongated opening defines an inducer section and a blade receipt section to retain a hollow fan blade section. The blade receipt section retains each of the hollow fan blade sections adjacent each inducer section. The inducer sections are cast directly into the fan hub which minimizes leakage between each fan blade section and each of the respective inducer sections to minimize airflow leakage and increase engine efficiency.

This invention was made with government support under Contract No.:F33657-03-C-2044. The government therefore has certain rights in thisinvention.

BACKGROUND OF THE INVENTION

The present invention relates to a tip turbine engine, and moreparticularly to a fan-turbine rotor assembly which includes an inducerformed therein.

An aircraft gas turbine engine of the conventional turbofan typegenerally includes a forward bypass fan, a compressor, a combustor, andan aft turbine all located along a common longitudinal axis. Acompressor and a turbine of the engine are interconnected by a shaft.The compressor is rotatably driven to compress air entering thecombustor to a relatively high pressure. This pressurized air is thenmixed with fuel in a combustor and ignited to form a high energy gasstream. The gas stream flows axially aft to rotatably drive the turbinewhich rotatably drives the compressor through the shaft. The gas streamis also responsible for rotating the bypass fan. In some instances,there are multiple shafts or spools. In such instances, there is aseparate turbine connected to a separate corresponding compressorthrough each shaft. In most instances, the lowest pressure turbine willdrive the bypass fan.

Although highly efficient, conventional turbofan engines operate in anaxial flow relationship. The axial flow relationship results in arelatively complicated elongated engine structure of considerablelongitudinal length relative to the engine diameter. This elongatedshape may complicate or prevent packaging of the engine into particularapplications.

A recent development in gas turbine engines is the tip turbine engine.Tip turbine engines locate an axial compressor forward of a bypass fanwhich includes hollow fan blades that receive airflow from the axialcompressor therethrough such that the hollow fan blades operate as acentrifugal compressor. Compressed core airflow from the hollow fanblades is mixed with fuel in an annular combustor and ignited to form ahigh energy gas stream which drives the turbine integrated onto the tipsof the hollow bypass fan blades for rotation therewith as generallydisclosed in U.S. Patent Application Publication Nos.: 20030192303;20030192304; and 20040025490.

The tip turbine engine provides a thrust to weight ratio equivalent toconventional turbofan engines of the same class within a package ofsignificantly shorter length.

One significant rotational component of a tip turbine engine is thefan-turbine rotor assembly. The fan-turbine rotor assembly includes amultitude of components which rotate at relatively high speeds togenerate bypass airflow while communicating a core airflow through eachof the multitude of hollow fan blades. A large percentage of the expenseassociated with a tip turbine engine is the manufacture of thefan-turbine rotor assembly and the integration of the inducer with thefan hub.

Accordingly, it is desirable to provide an inducer arrangement for afan-turbine rotor assembly, which is relatively inexpensive tomanufacture yet provides a high degree of reliability.

SUMMARY OF THE INVENTION

The fan-turbine rotor assembly for a tip turbine engine according to thepresent invention includes a fan hub which has an outer peripheryscalloped by a multitude of elongated openings which extend into a fanhub web. Each elongated opening defines an inducer section and a bladereceipt section to retain a hollow fan blade section. The blade receiptsection retains each of the hollow fan blade sections adjacent eachinducer section. An inner fan blade mount is located adjacent an inducerexhaust section to communicate a core airflow communication path fromwithin each inducer section into the core airflow passage within eachfan blade section.

The inducer is cast directly into the fan hub which minimizes leakagebetween each fan blade section and each inducer section to provideincreased engine efficiency. Manufacturing and assembly is also readilyfacilitated.

The present invention therefore provides an inducer arrangement for afan-turbine rotor assembly which is relatively inexpensive tomanufacture yet provides a high degree of reliability.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of the currently preferred embodiment. The drawings thataccompany the detailed description can be briefly described as follows:

FIG. 1 is a partial sectional perspective view of a tip turbine engine;

FIG. 2 is a longitudinal sectional view of a tip turbine engine along anengine centerline;

FIG. 3 is an exploded view of a fan-turbine rotor assembly;

FIG. 4 is an assembled view of a fan-turbine rotor assembly;

FIG. 5A is an expanded radial sectional view of an inducer section;

FIG. 5B is a sequential sectional view of the fan hub illustrating theinducer sections therewith;

FIG. 6 is a schematic view of airflow through the last stage of an axialcompressor and into the inducer;

FIG. 7A is an expanded phantom perspective view of a fan blade mountedto a hub of a fan-turbine rotor assembly;

FIG. 7B is an expanded partially sectioned perspective view of a fanblade mounted to a hub of a fan-turbine rotor assembly; and

FIG. 7C is an expanded partially sectioned perspective view of adiffuser section of a fan blade.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 illustrates a general perspective partial sectional view of a tipturbine engine type gas turbine engine 10. The engine 10 includes anouter nacelle 12, a rotationally fixed static outer support structure 14and a rotationally fixed static inner support structure 16. A multitudeof fan inlet guide vanes 18 are mounted between the static outer supportstructure 14 and the static inner support structure 16. Each inlet guidevane preferably includes a variable trailing edge 18A.

A nose cone 20 is preferably located along the engine centerline A tosmoothly direct airflow into an axial compressor 22 adjacent thereto.The axial compressor 22 is mounted about the engine centerline A behindthe nose cone 20.

A fan-turbine rotor assembly 24 is mounted for rotation about the enginecenterline A aft of the axial compressor 22. The fan-turbine rotorassembly 24 includes a multitude of hollow fan blades 28 to provideinternal, centrifugal compression of the compressed airflow from theaxial compressor 22 for distribution to an annular combustor 30 locatedwithin the rotationally fixed static outer support structure 14.

A turbine 32 includes a multitude of tip turbine blades 34 (two stagesshown) which rotatably drive the hollow fan blades 28 relative amultitude of tip turbine stators 36 which extend radially inwardly fromthe static outer support structure 14. The annular combustor 30 isaxially forward of the turbine 32 and communicates with the turbine 32.

Referring to FIG. 2, the rotationally fixed static inner supportstructure 16 includes a splitter 40, a static inner support housing 42and an static outer support housing 44 located coaxial to said enginecenterline A.

The axial compressor 22 includes the axial compressor rotor 46 fromwhich a plurality of compressor blades 52 extend radially outwardly anda compressor case 50 fixedly mounted to the splitter 40. A plurality ofcompressor vanes 54 extend radially inwardly from the compressor case 50between stages of the compressor blades 52. The compressor blades 52 andcompressor vanes 54 are arranged circumferentially about the axialcompressor rotor 46 in stages (three stages of compressor blades 52 andcompressor vanes 54 are shown in this example). The axial compressorrotor 46 is mounted for rotation upon the static inner support housing42 through a forward bearing assembly 68 and an aft bearing assembly 62.

The fan-turbine rotor assembly 24 includes a fan hub 64 that supports amultitude of the hollow fan blades 28. Each fan blade 28 includes aninducer section 66, a hollow fan blade section 72 and a diffuser section74. The inducer section 66 receives airflow from the axial compressor 22generally parallel to the engine centerline A and turns the airflow froman axial airflow direction toward a radial airflow direction. Theairflow is radially communicated through a core airflow passage 80within the fan blade section 72 where the airflow is centrifugallycompressed. From the core airflow passage 80, the airflow is turned anddiffused toward an axial airflow direction toward the annular combustor30. Preferably the airflow is diffused axially forward in the engine 10,however, the airflow may alternatively be communicated in anotherdirection.

A gearbox assembly 90 aft of the fan-turbine rotor assembly 24 providesa speed increase between the fan-turbine rotor assembly 24 and the axialcompressor 22. Alternatively, the gearbox assembly 90 could provide aspeed decrease between the fan-turbine rotor assembly 24 and the axialcompressor rotor 46. The gearbox assembly 90 is mounted for rotationbetween the static inner support housing 42 and the static outer supporthousing 44. The gearbox assembly 90 includes a sun gear shaft 92 whichrotates with the axial compressor 22 and a planet carrier 94 whichrotates with the fan-turbine rotor assembly 24 to provide a speeddifferential therebetween. The gearbox assembly 90 is preferably aplanetary gearbox that provides co-rotating or counter-rotatingrotational engagement between the fan-turbine rotor assembly 24 and anaxial compressor rotor 46. The gearbox assembly 90 is mounted forrotation between the sun gear shaft 92 and the static outer supporthousing 44 through a forward bearing 96 and a rear bearing 98. Theforward bearing 96 and the rear bearing 98 are both tapered rollerbearings and both handle radial loads. The forward bearing 96 handlesthe aft axial loads while the rear bearing 98 handles the forward axialloads. The sun gear shaft 92 is rotationally engaged with the axialcompressor rotor 46 at a splined interconnection 100 or the like.

In operation, air enters the axial compressor 22, where it is compressedby the three stages of the compressor blades 52 and compressor vanes 54.The compressed air from the axial compressor 22 enters the inducersection 66 in a direction generally parallel to the engine centerline Aand is turned by the inducer section 66 radially outwardly through thecore airflow passage 80 of the hollow fan blades 28. The airflow isfurther compressed centrifugally in the hollow fan blades 28 by rotationof the hollow fan blades 28. From the core airflow passage 80, theairflow is turned and diffused axially forward in the engine 10 into theannular combustor 30. The compressed core airflow from the hollow fanblades 28 is mixed with fuel in the annular combustor 30 and ignited toform a high-energy gas stream. The high-energy gas stream is expandedover the multitude of tip turbine blades 34 mounted about the outerperiphery of the fan-turbine rotor assembly 24 to drive the fan-turbinerotor assembly 24, which in turn drives the axial compressor 22 throughthe gearbox assembly 90. Concurrent therewith, the fan-turbine rotorassembly 24 discharges fan bypass air axially aft to merge with the coreairflow from the turbine 32 in an exhaust case 106. A multitude of exitguide vanes 108 are located between the static outer support housing 44and the rotationally fixed static outer support structure 14 to guidethe combined airflow out of the engine 10 to provide forward thrust. Anexhaust mixer 110 mixes the airflow from the turbine blades 34 with thebypass airflow through the fan blades 28.

Referring to FIG. 3, the fan-turbine rotor assembly 24 is illustrated inan exploded view. The fan hub 64 is the primary structural support ofthe fan-turbine rotor assembly 24 (FIG. 4). The fan hub 64 is preferablyforged and then milled to provide the desired geometry. The fan hub 64defines a bore 111 and an outer periphery 112. The outer periphery 112is preferably scalloped by a multitude of elongated openings 111. Thefan hub 64 is the primary structural support of the fan-turbine rotorassembly 24. The fan hub 64 supports the multitude of fan blades 28, adiffuser 114, and the turbine 32. The diffuser 114 defines a diffusersurface 119 formed about the outer periphery of the fan blade sections72 to provide structural support to the outer tips of the fan bladesections 72 and to turn and diffuse the airflow from the radial coreairflow passage 80 (FIG. 3) toward an axial airflow direction. Theturbine 32 is mounted to the diffuser surface 119 as one or more turbinering rotors 118 a, 118 b which may include a multitude of turbine bladeclusters.

Referring to FIG. 4, the fan hub 64 itself forms the multitude ofinducer sections 66. Each inducer section 66 formed by the fan hub 64 isessentially a conduit that defines an inducer passage 118 between aninducer inlet section 120 and an inducer exit section 128 FIGS. 5A, 5B).

Referring to FIGS. 5A and 5B, the inducer sections 66 together form theinducer 116 of the fan-turbine rotor assembly 24. The inducer inletsection 120 of each inducer passage 118 extends forward of the fan hub64 and is canted toward a rotational direction of the fan hub 64 suchthat inducer inlet 120 operates as an air scoop during rotation of thefan-turbine rotor assembly 24. Each inducer passage 118 providesseparate airflow communication to each core airflow passage 80 when eachfan blade section 72 is mounted within each elongated opening 114.Preferably, each fan blade section 72 includes an attached diffusersection 74 such that the diffuser surface 119 is formed when thefan-turbine rotor assembly 24 is assembled.

FIG. 6 schematically illustrates the relationship of the angle of thelast stage of the compressor rotor blade 52 (one shown) and the laststage of the compressor vanes 54 in the three stage axial compressor 22(FIG. 2) prior to communication of the airflow from the axial compressor22 into the inducer sections 66 in the engine 10. Referring to thecompressor blade velocity triangle Bt, the compressor rotor blade 52 isangled relative to the engine centerline A to provide an angle of arelative velocity vector, Vr1. The velocity of the counter-rotatingcompressor blade 52 gives a blade velocity vector, Vb1. The resultantvector, indicating the resultant core airflow from the compressor blade52, is the absolute velocity vector, Val.

Referring to the vane velocity vector St, a stator leading edge 541 ofthe compressor stator 54 is angled to correspond with the absolutevelocity vector, Va1 from the compressor rotor blade 52 to efficientlyreceive and compress the core airflow from the compressor blade 52. Thevane trailing edge 54 t is angled relative to the engine centerline A tocompress and redirect the airflow toward the inducer section 66 (oneshown) as the inducer 116 rotates relative thereto at a vane absolutevelocity vector, Va1.

The inducer inlet 120 of the inducer section 66 is angled to efficientlyreceive the core airflow from the vane trailing edge 54 t which flowstoward the inducer section 66 at the absolute velocity vector, Va1 fromthe vane 54. The velocity of the inducer section 66 gives an inducervelocity vector, Vb1. Referring to the inducer velocity triangle It, theangle of the inducer 66 is selected such that the sum of the inducerrelative velocity vector Vr1 and the inducer velocity vector Vb1 matchthe angle of the core airflow incoming from the compressor vane trailingedge 54 t (absolute velocity vector, Val).

It should be understood that the specific angles will depend on avariety of factors, including anticipated blade velocities and thedesign choices made in the earlier stages of the compressor blades 52and compressor vanes 54 to provide a length sufficient to turn the coreairflow from axial flow to radial flow while decreasing the overalllength of the engine 10. It should be understood that the axialcompressor 22 may alternatively counter-rotate relative to inducer 116as disclosed in co-pending application ______ entitled “COUNTER-ROTATINGGEARBOX FOR TIP TURBINE ENGINE,” which is assigned to the assignee ofthe present invention and which is hereby incorporated by reference inits entirety.

Referring to FIG. 7A, the fan hub 64 retains each hollow fan bladesection 72 through a blade receipt section 122. The blade receiptsection 122 preferably forms an axial semi-cylindrical opening formedalong the axial length of the elongated openings 111. It should beunderstood that other retention structures such as a dove-tail,fir-tree, or bulb-type engagement structure will likewise be usable withthe present invention.

Each hollow fan blade section 72 includes a fan blade mount section 124that corresponds with the blade receipt section 122 to retain the hollowfan blade section 72 within the fan hub 64. The fan blade mount 124preferably includes a semi-cylindrical portion to radially retain thefan blade 28.

Referring to FIG. 7B, the inner fan blade mount 124 is preferablyuni-directionally mounted into the blade receipt section 122 such asfrom the rear face of the fan hub 64. The fan blade mount section 124engages the blade receipt section 122 during operation of thefan-turbine rotor assembly 24 to provide a directional locktherebetween. That is, the inner fan blade mount 124 and the bladereceipt section 122 may be frustoconical or axially non-symmetrical suchthat the forward segments form a smaller perimeter than the rear segmentto provide a wedged engagement therebetween when assembled.

Each inducer section 66 within the fan hub 64 receives core airflowcommunication from the inducer passages 118 into the core airflowpassage 80 and turns and diffuses the airflow through each diffusersection 74 of the diffuser 114 (also illustrated in FIG. 7C).

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

The foregoing description is exemplary rather than defined by thelimitations within. Many modifications and variations of the presentinvention are possible in light of the above teachings. The preferredembodiments of this invention have been disclosed, however, one ofordinary skill in the art would recognize that certain modificationswould come within the scope of this invention. It is, therefore, to beunderstood that within the scope of the appended claims, the inventionmay be practiced otherwise than as specifically described. For thatreason the following claims should be studied to determine the truescope and content of this invention.

1. A fan hub assembly for a tip turbine engine comprising: a fan hubdefining a hub axis of rotation, said fan hub defining a multitude ofelongated openings through an outer periphery of said fan hub; a bladereceipt section defined by each of said elongated openings; and aninducer section defined within each of said elongated openings to turnan airflow from a generally axial direction to a generally radialdirection.
 2. The fan hub assembly as recited in claim 1, wherein saidinducer section is cast within said fan hub.
 3. The fan hub assembly asrecited in claim 1, further comprising a multitude of fan blades, eachof said multitude of fan blade receivable within each of said bladereceipt sections to receive an airflow through a core airflow passagedefined within each of said fan blades.
 4. The fan hub assembly asrecited in claim 3, wherein each of said multitude of fan blades includea fan blade mount section receivable within each of said blade receiptsections.
 5. The fan hub assembly as recited in claim 4, wherein each ofsaid fan blade mount sections are semi-cylindrical to radially lock saidfan blade sections within said fan hub.
 6. A fan-turbine rotor assemblyfor a tip turbine engine comprising: a fan hub defining a hub axis ofrotation, said fan hub defining a multitude of elongated openingsthrough an outer periphery of said fan hub; an inducer defined by eachof said elongated openings to turn an airflow from a generally axialdirection to a generally radial direction; a blade receipt sectiondefined by each of said elongated openings; a multitude of fan bladesections which each define a core airflow passage therethrough; and afan blade mount section extending from each of said multitude of fanblade sections, each of said fan blade mount sections receivable withinone of said multitude of blade receipt sections for retention therein tocommunicate said airflow from said inducer to each of said multitude ofcore airflow passages.
 7. The fan-turbine rotor assembly as recited inclaim 6, further comprising a diffuser about said multitude of fan bladesections, said diffuser in communication with each of said multitude ofcore airflow passages to turn said airflow from said radial direction toa second axial airflow direction.
 8. The fan-turbine rotor assembly asrecited in claim 7, further comprising a turbine which extends from saiddiffuser.
 9. The fan-turbine rotor assembly as recited in claim 8,wherein said turbine includes a first row of shrouded turbine blades anda second row of shrouded turbine blades.